JP2684936B2 - Gas turbine and gas turbine blade - Google Patents
Gas turbine and gas turbine bladeInfo
- Publication number
- JP2684936B2 JP2684936B2 JP4249933A JP24993392A JP2684936B2 JP 2684936 B2 JP2684936 B2 JP 2684936B2 JP 4249933 A JP4249933 A JP 4249933A JP 24993392 A JP24993392 A JP 24993392A JP 2684936 B2 JP2684936 B2 JP 2684936B2
- Authority
- JP
- Japan
- Prior art keywords
- blade
- edge side
- arc
- trailing edge
- gas turbine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Lifetime
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2212—Improvement of heat transfer by creating turbulence
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Description
【０００１】[0001]
【産業上の利用分野】本発明は、内部より冷却媒体によ
り冷却されるように形成されている翼を備えたガスター
ビン及びその翼の改良に関するものである。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine provided with blades formed so as to be cooled from inside by a cooling medium, and an improvement of the blades.
【０００２】[0002]
【従来の技術】ガスタービンエンジンの性能を向上させ
るために、最近益々の燃焼ガスの温度を上げることが行
われ、ガスタービンの翼は熱的に非常に苛酷な環境下に
おいて作動している。2. Description of the Related Art Recently, in order to improve the performance of gas turbine engines, the temperature of combustion gas has been increased more and more, and the blades of gas turbines are operated in a very thermally severe environment.
【０００３】従ってこれらの翼は何らかの冷却手段によ
り充分な冷却が行われなければならない。Therefore, these blades must be sufficiently cooled by some cooling means.
【０００４】一般にこの種タービン翼の冷却は、圧縮さ
れた燃焼用空気の一部を翼内部の空洞部に流通させるこ
とにより冷却するようにしたものが広く採用されてい
る。その代表的な翼の冷却例は、例えば特開昭2−24190
2 号公報にも示されている。Generally, this kind of turbine blade is widely cooled by cooling a part of compressed combustion air by flowing it into a cavity inside the blade. A typical example of cooling the blade is disclosed in, for example, Japanese Patent Laid-Open No.
It is also shown in Publication No. 2.
【０００５】一方この種翼の形状は、翼型の中心である
矢高線が円弧，放物線の弧、或いは滑らかに変わる他の
曲線の一部として与えられ、この矢高線に沿って翼形が
設定されている。この場合厚みは翼の前縁から後縁に向
かうにしたがい徐々に厚くして最大厚みをとり、その後
後縁側に向かって薄くなるように形成されている。On the other hand, the shape of this seed wing is given as a part of the arrow height line which is the center of the airfoil as a circular arc, a parabolic arc, or another curve that smoothly changes, and the airfoil shape is set along this arrow height line. Has been done. In this case, the thickness is gradually increased from the leading edge to the trailing edge of the blade to reach the maximum thickness, and then is reduced toward the trailing edge side.
【０００６】[0006]
【発明が解決しようとする課題】ガスタービン翼はこの
ように形成され、かつ前述したようにその内部から冷却
されるわけであるが、ガスタービンの場合この冷却に用
いられる空気は燃焼用空気の一部を取り出して用いてい
るのが普通である。このため冷却空気の消費量が多いと
燃焼用の空気が制約されることになり、ガスタービン高
温化によるサイクル上のメリットを損なうことになる。
このため、翼の冷却に用いられる冷却空気の量はできる
だけ少量であることが望ましい。The gas turbine blade is formed in this way and is cooled from the inside as described above. In the case of a gas turbine, the air used for cooling is the combustion air. It is normal to take out a part and use it. For this reason, if the consumption of cooling air is large, the combustion air is restricted, and the cycle merit due to the high temperature of the gas turbine is impaired.
Therefore, it is desirable that the amount of cooling air used for cooling the blade is as small as possible.
【０００７】本発明はこれに鑑みなされたものでその目
的とするところは、使用冷却空気量少なくして翼の冷却
が充分有効に行われるガスタービンの翼及び充分高温化
が可能なこの種ガスタービンを提供するにある。The present invention has been made in view of the above, and an object of the present invention is to provide a blade of a gas turbine in which the amount of cooling air used is reduced and the blade is sufficiently cooled, and a gas of this kind capable of sufficiently increasing the temperature. To provide a turbine.
【０００８】[0008]
【課題を解決するための手段】すなわち本発明は、翼厚
が前縁側より後縁側にかけて徐々に小さくなるように形
成され、かつ内部に冷却媒体流通路を有している翼にお
いて、翼前縁側の横断面形状を円弧状にするとともに、
翼の最大厚み部分がこの円弧の中心部分に位置するよう
に形成し所期の目的を達成するようにしたものである。That is, the present invention relates to a blade having a cooling medium flow passage formed therein such that the blade thickness gradually decreases from the leading edge side toward the trailing edge side, and the blade leading edge side is provided. In addition to making the cross-sectional shape of the arc circular,
The blade is formed so that the maximum thickness portion thereof is located at the central portion of this arc so as to achieve the intended purpose.
【０００９】[0009]
【作用】すなわちこのように形成された翼であると、主
流ガスが、翼の腹側と滑らかに接続した前縁円弧上の一
端点部分および翼の背側と滑らかに接続した前縁円弧上
の一端点部分に沿って流れるため、前縁部の急激な高温
ガスの加速が抑制される。すなわち翼表面の高温ガスの
流速を小さくすることができるので、ガス側の熱伝達率
が下がり、したがって翼内部を流通させる冷却空気の量
を少なくすることができるのである。In other words, with the blade thus formed, the mainstream gas flows on the leading edge arc smoothly connected to the ventral side of the blade and on the leading edge arc smoothly connected to the back side of the blade. Since it flows along the one end point of the, the rapid acceleration of the high temperature gas at the front edge is suppressed. That is, since the flow velocity of the hot gas on the blade surface can be reduced, the heat transfer coefficient on the gas side is reduced, and therefore the amount of cooling air flowing inside the blade can be reduced.
【００１０】[0010]
【実施例】以下図示した実施例に基づいて本発明を詳細
に説明する。DESCRIPTION OF THE PREFERRED EMBODIMENTS The present invention will be described below in detail with reference to the illustrated embodiments.
【００１１】図３にはそのガスタービンが一部断面して
示されている。図中１はローター、２はステーターであ
る。ローター１は主として回転軸３、この回転軸に配置
された動翼４および圧縮機５の動翼とを備え、又ステー
ター２は主としてケーシング７、このケーシングに支持
され前記動翼に対向するように配置されている燃焼器
８、それに燃焼器のノズルの役をなす静翼９を備えてい
る。FIG. 3 shows the gas turbine in a partial cross section. In the figure, 1 is a rotor and 2 is a stator. The rotor 1 mainly comprises a rotating shaft 3, moving blades 4 arranged on this rotating shaft, and moving blades of the compressor 5, and the stator 2 mainly comprises a casing 7, which is supported by the casing and faces the moving blades. It has a combustor 8 arranged therein and a vane 9 acting as a nozzle of the combustor.
【００１２】このように構成されているガスタービンの
概略動作は、まず圧縮機５からの圧縮空気と燃料が燃焼
器８に与えられ、この燃焼器内でこれら燃料が燃焼し高
温ガスを発生する。そして発生した高温ガスは静翼９を
介して動翼４に吹きつけられ動翼を介してローターを駆
動する。The general operation of the gas turbine thus constructed is that compressed air from the compressor 5 and fuel are first supplied to the combustor 8, and the fuel is combusted in the combustor to generate high temperature gas. . Then, the generated high-temperature gas is blown to the moving blade 4 via the stationary blade 9 and drives the rotor via the moving blade.
【００１３】この場合高温ガス中にさらされている動翼
４や静翼９は冷却する必要があり、その冷却媒体には前
記圧縮機５の圧縮空気の一部が用いられている。In this case, the moving blades 4 and the stationary blades 9 exposed to the high temperature gas need to be cooled, and a part of the compressed air of the compressor 5 is used as the cooling medium.
【００１４】図４はその静翼の冷却の一例を示したもの
である。この図は静翼９と動翼４の部分、すなわち段落
部を示したもので、静翼９が断面して示されている。FIG. 4 shows an example of cooling the stationary blade. This figure shows a part of the stationary blade 9 and the moving blade 4, that is, a paragraph, and the stationary blade 9 is shown in cross section.
【００１５】静翼９は外周壁１０と内周壁１１の間にこ
れらの壁に固着されて設けられている。内周壁１１には
回転子３との間隙に上流側と下流側とを隔てる仕切装置
１２が設けられている。冷却空気は外周壁１０に設けら
れた冷却空気導入孔１０ａより、冷却空気供給源、すな
わち圧縮機５（図３参照）から翼内の空気冷却室９ｆに
導かれる。The stationary blade 9 is provided between the outer peripheral wall 10 and the inner peripheral wall 11 so as to be fixed to these walls. The inner peripheral wall 11 is provided with a partitioning device 12 that separates the upstream side and the downstream side in a gap with the rotor 3. The cooling air is introduced from the cooling air supply source, that is, the compressor 5 (see FIG. 3), into the air cooling chamber 9f in the blade through the cooling air introduction hole 10a provided in the outer peripheral wall 10.
【００１６】冷却後の冷却空気は内周壁に設けられた排
出孔１１ａより排出され、やがてはガスパス路に排出さ
れる。After cooling, the cooling air is discharged from the discharge hole 11a provided in the inner peripheral wall, and eventually to the gas path.
【００１７】尚図中線矢印は冷却空気の流れを示し、枠
どり矢印は高温ガス、すなわち主流作動ガスの流れを示
している。The arrows in the figure show the flow of cooling air, and the framed arrows show the flow of hot gas, that is, the mainstream working gas.
【００１８】静翼９はこのように内部から冷却されるわ
けであるが、この静翼は特に次のような形状に形成され
ている。図１にはその静翼が横断面で示され、図２には
その線図が示されている。The stationary blade 9 is cooled from the inside in this way, and the stationary blade is formed in the following shape. The stator vane is shown in cross section in FIG. 1 and its diagram in FIG.
【００１９】これらの図において、９ａはその前縁、９
ｂは翼背側部、９ｃは翼腹側部、９ｄは後縁部である。
図中９ｆが前述した空気冷却室である。この空気冷却室
は外被および隔壁により３個の空洞、すなわち空気冷却
室９ｆ１，９ｆ２，９ｆ３に分割されている。この場合
翼の前部の空気冷却室９ｆ１，９ｆ２には熱変換を良好
にするためにフィン９ｈが設けられ、翼後縁部の空気冷
却室９ｆ３にはピンフィン９ｇが配設されている。なお
この冷却構造は対流冷却や他の冷却手段であっても構わ
ない。In these figures, 9a is its leading edge, 9a
Reference numeral b is a blade back side portion, 9c is a blade ventral side portion, and 9d is a trailing edge portion.
In the figure, 9f is the above-mentioned air cooling chamber. This air cooling chamber is divided into three cavities, that is, air cooling chambers 9f1, 9f2, 9f3 by a jacket and partition walls. In this case, fins 9h are provided in the air cooling chambers 9f1 and 9f2 at the front of the blade to improve heat conversion, and pin fins 9g are provided in the air cooling chamber 9f3 at the trailing edge of the blade. The cooling structure may be convection cooling or other cooling means.
【００２０】翼の冷却構造は以上のようであるが、ここ
で最も重要なことはこの翼自体の全体の形は次のように
形成されていることである。すなわち前縁１側の横断面
形状が円弧状（直径Ｄ２）に形成されており、そしてこ
の円弧の最大太り部分より後縁側に向かうにしたがいそ
の翼厚みを徐々に小さくなるように形成されている。な
おこの場合、円弧の最大太り部といっても次第に厚みが
小さく部分と結合されるので、正確には多少ずれる、す
なわち図中Ｓ１，Ｐ１となる。The cooling structure of the blade is as described above, but the most important thing here is that the entire shape of the blade itself is formed as follows. That is, the cross-sectional shape of the leading edge 1 side is formed in an arc shape (diameter D2), and the blade thickness is gradually reduced toward the trailing edge side from the maximum thickened portion of this arc. . In this case, since the maximum thickness of the circular arc is gradually reduced, the maximum thickness of the circular arc is combined with that of the circular arc.
【００２１】また後縁に向かうにしたがい厚みが薄くな
ると云っても翼後縁部は機械的な強度の問題もあり、あ
まり薄くはできないので後縁には小さな円弧部が設けら
れている。換言すればこの翼形状（横断面）は概略的に
は人玉状に形成されるということである。Further, even if it is said that the thickness becomes thinner toward the trailing edge, there is a problem in mechanical strength at the trailing edge portion of the blade, and since it cannot be made too thin, a small arc portion is provided at the trailing edge. In other words, this wing shape (transverse cross section) is roughly formed into a human ball shape.
【００２２】静翼９はこのように形成されているわけで
あるが、次にこの翼の作用について従来の翼と比較しな
がら説明する。The stationary blade 9 is formed in this way. Next, the operation of this blade will be described in comparison with a conventional blade.
【００２３】図５，図６は夫々翼を線図で表わし並設し
たもので、Ｎ1 が従来の翼、Ｎ2 が本発明の翼である。
Ｎ1 翼，Ｎ2 翼は同一性能のもので、回転軸の周囲に設
けられる数も同じ、すなわちピッチＳ１は同一である。FIG. 5 and FIG. 6 show the blades arranged side by side in a diagram, where N 1 is the conventional blade and N 2 is the blade of the present invention.
The N 1 blade and the N 2 blade have the same performance, and the numbers provided around the rotating shaft are the same, that is, the pitch S1 is the same.
【００２４】体格的には、最大コード長はＮ1 翼の方が
Ｎ2 翼より大きい、すなわちＣ１＞Ｃ２となる。この比
較の値が表１に示されている。Physically, the maximum cord length of the N 1 blade is larger than that of the N 2 blade, that is, C1> C2. The values for this comparison are shown in Table 1.
【００２５】[0025]
【表１】 [Table 1]
【００２６】すなわち表面積はコード長が短いＮ2 翼が
Ｎ1 翼の９１％、前縁径はＮ２翼がＮ1 翼の１.６ 倍、
段面積はＮ2 翼がＮ1 翼の８９％である。That is, the surface area of the N 2 blade having a short cord length is 91% of that of the N 1 blade, and the leading edge diameter of the N 2 blade is 1.6 times that of the N 1 blade.
The step area is 89% for N 2 blades and N 1 blades.
【００２７】この２つの翼の全圧損失係数と、翼メタル
温度を材料の許容温度まで冷却するのに必要な空気流量
を求めたのが表２である。Table 2 shows the total pressure loss coefficient of these two blades and the air flow rate required for cooling the blade metal temperature to the allowable temperature of the material.
【００２８】[0028]
【表２】 [Table 2]
【００２９】全圧損失係数はＮ1 翼，Ｎ2 翼とも変わら
ず同じ翼じ値であった。又実験の結果では冷却空気の消
費量はＮ2 翼が従来のＮ1 翼より８％少なかった。これ
は翼表面積が減少したことと、前縁径が円弧状をなし大
きいことによる。次に前縁径が大きくなれば冷却空気の
消費量が低減することについて説明する。The total pressure loss coefficient was the same for both N 1 blade and N 2 blade. Also, the result of the experiment shows that the consumption of cooling air by the N 2 blade is 8% less than that of the conventional N 1 blade. This is because the blade surface area was reduced and the leading edge diameter was arcuate and large. Next, it will be explained that the consumption amount of the cooling air decreases as the leading edge diameter increases.
【００３０】翼前縁ガス側の熱伝達率αg はThe heat transfer coefficient α g on the blade leading edge gas side is
【００３１】[0031]
【数１】 (Equation 1)
【００３２】 ｋ1，ｋ2；定数 Ｒｅ；レイノルズ数 Ｐｒ；プラントル数 Ｖ；ガス流速 Ｄ；前縁径 λ；熱伝導率 ν；動粘性係数 で表される。ガス側の条件が同じであればK 1 , k 2 ; constant Re; Reynolds number Pr; Prandtl number V; gas flow rate D; leading edge diameter λ; thermal conductivity ν; If the conditions on the gas side are the same,
【００３３】[0033]
【数２】 (Equation 2)
【００３４】となる。従ってＮ2 翼とＮ1 翼のガス側熱
伝達率の比は## EQU1 ## Therefore, the ratio of the heat transfer coefficient on the gas side between the N 2 blade and the N 1 blade is
【００３５】[0035]
【数３】 (Equation 3)
【００３６】となり、Ｎ2 翼は従来のＮ1 翼よりも２１
％ガス側の熱伝達率が下がるので高温ガスから翼への伝
熱量も減少し、その結果として少ない冷却空気で翼前縁
を冷却することができるということである。Therefore, the N 2 blade is 21 more than the conventional N 1 blade.
Since the heat transfer coefficient on the% gas side is reduced, the amount of heat transfer from the hot gas to the blade is also reduced, and as a result, the blade leading edge can be cooled with less cooling air.
【００３７】冷却空気の消費量が低減すれば単にガスタ
ービンのサイクル効率が向上するだけでなく冷却空気と
主流ガスの混合損失も減少するので、ガスタービンの性
能改善効果は大きい。また翼断面積も１１％減少してい
るので、翼材料費が低減するという効果もある。If the consumption of the cooling air is reduced, not only the cycle efficiency of the gas turbine is improved but also the mixing loss of the cooling air and the mainstream gas is reduced, so that the performance of the gas turbine is greatly improved. Further, since the blade cross-sectional area is reduced by 11%, the blade material cost is also reduced.
【００３８】本発明の他の実施例を図６を用いて説明す
る。Another embodiment of the present invention will be described with reference to FIG.
【００３９】図６において、図１と同一記号であれば、
図１と同一構成，機能を有する。In FIG. 6, if the same symbols as in FIG.
It has the same configuration and function as in FIG.
【００４０】Ｎ3 翼は前縁径がＤ3（＞Ｄ2＞Ｄ1）であ
り、翼厚み分布は図２のＮ2翼と同じように翼前縁から
翼後縁に至るまで単調に減少している。The leading edge diameter of the N 3 blade is D 3 (> D 2 > D 1 ), and the blade thickness distribution is monotonous from the leading edge to the trailing edge of the blade as in the N 2 blade of FIG. is decreasing.
【００４１】翼最大コード長は従来のＮ1 翼と同じ値と
なっているのが、ピッチＳ3 は、従来よりも広くしてあ
る。この形状を比較して表３に示す。The maximum blade chord length is the same as that of the conventional N 1 blade, but the pitch S 3 is wider than that of the conventional blade. The shapes are compared and shown in Table 3.
【００４２】[0042]
【表３】 [Table 3]
【００４３】翼１枚当りの表面積はＮ3 翼，Ｎ1 翼同一
である。しかしＮ3 翼は翼枚数をＮ 1 翼よりも２４％削
減しているので、翼列全体で考えると、全表面積は２４
％減少したことになる。このＮ3 ，Ｎ1 翼の全圧損失係
数と冷却空気の消費量を求め表４に示した。The surface area per blade is NThreeWings, N1Same wing
It is. But NThreeThe number of wings is N 124% less than wings
Since it has decreased, the total surface area is 24 when considering the entire cascade.
It means that it has decreased by%. This NThree, N1Wing total pressure loss
The number and the consumption of cooling air were determined and are shown in Table 4.
【００４４】[0044]
【表４】 [Table 4]
【００４５】全圧損失係数はＮ3 翼が従来のＮ1 翼の７
７％であり、２３％減少している。これは、Ｎ3 翼の後
縁厚みは、Ｎ1 翼と同一であるが、翼枚数が２４％減少
したことにより、ピッチＳ3 が従来翼のピッチＳ1 の
１.３１ 倍となり相対的な後縁厚み（後縁厚／ピッチ）
が小さくなり、後縁損失が低減したことによる。As for the total pressure loss coefficient, N 3 blade is 7 times that of the conventional N 1 blade.
7% and 23% decrease. This edge thickness after N 3 wing is identical to the N 1 wing by wing number was reduced 24%, the pitch S 3 is relative becomes 1.31 times the pitch S 1 in the conventional blade Rear edge thickness (rear edge thickness / pitch)
Is smaller and the trailing edge loss is reduced.
【００４６】冷却空気の消費量は、翼全表面積が２４％
減少したことと、前縁径Ｄ3 が従来翼Ｎ1 の２.１ 倍に
なったことにより、２０％減少している。先に説明した
ガス側の熱伝達率を比較するとCooling air consumption is 24% of the total blade surface area.
The decrease is 20% due to the decrease and the leading edge diameter D 3 being 2.1 times as large as that of the conventional blade N 1 . Comparing the heat transfer coefficients on the gas side described above,
【００４７】[0047]
【数４】 (Equation 4)
【００４８】と、３１％も減少している。Then, it has decreased by 31%.
【００４９】このようにＮ3 翼は従来翼Ｎ1 にくらべ冷
却空気の消費量だけでなく全圧損失係数も小さくなって
いるので、ガスタービン効率の向上効果は大きい。As described above, the N 3 blade has a smaller total air pressure consumption coefficient and a smaller total pressure loss coefficient than the conventional blade N 1 , so that the effect of improving the gas turbine efficiency is great.
【００５０】次に前縁と翼背側及び翼腹側とを接続する
２つの端点について図７を用いて説明する。Next, two end points that connect the leading edge to the wing back side and the wing ventral side will be described with reference to FIG.
【００５１】この図においてＮ3 翼の背側の端点Ｓ
3 は、腹側の端点Ｐ3 と前縁１の中心Ｏを結ぶ直線より
も下流側（後縁側）に位置している。この関係は先に説
明したＮ2 翼でも保たれている。Ｎ4 翼は翼背側２と前
縁１の接続点Ｓ4 が、腹側の接続点Ｐ3 と前縁円弧中心
Ｏを結ぶ直線よりも上流側に位置している。この２つの
翼の空力性能を比較して図８に示す。In this figure, the dorsal end point S of the N 3 wing is shown.
3 is located on the downstream side (rear edge side) of the straight line connecting the ventral end point P 3 and the center O of the front edge 1. This relationship is maintained even in the N 2 wing described above. In the N 4 blade, a connection point S 4 between the blade back side 2 and the leading edge 1 is located upstream of a straight line connecting the ventral side connection point P 3 and the leading edge arc center O. A comparison of the aerodynamic performance of these two blades is shown in FIG.
【００５２】この図は翼面のマッハ数分布を示すもので
あり、横軸は翼の軸方向位置（前縁からの軸方向距離／
軸コード長）をとっている。This figure shows the Mach number distribution on the blade surface, where the horizontal axis is the axial position of the blade (axial distance from the leading edge /
Axis code length).
【００５３】翼背側の最大マッハ数はＮ4 翼の方がＮ3
翼よりも大きく、最大マッハ数位置翼後縁に至るまで急
激に減速しており、翼背側で流れの剥離が観察された。
その結果、全圧損失係数はＮ3 翼の１.９ 倍となり空力
損失の大きいことが分かった。これはＮ4 翼において、
円弧形状の前縁からＳ4 点で背側を形成する曲線に接続
する際の曲率半径の変化の大きいことが原因となってい
ることが分かった。この翼背側の急激な加速流れと減速
流れを防止するためには、図７に示したように腹側の端
点Ｐ3 と前縁円弧中心Ｏとを結ぶ直線よりも下流側に、
背側の端点Ｓ3を位置させればよいことを、翼間流れ計
算により見出した。The maximum Mach number on the back side of the wing is N 3 for the N 4 wing.
It was larger than the blade, and was rapidly decelerated to the trailing edge of the blade at the maximum Mach number position, and flow separation was observed on the back side of the blade.
As a result, it was found that the total pressure loss coefficient was 1.9 times that of the N 3 blade, resulting in a large aerodynamic loss. This is the N 4 wing,
It was found that the cause was a large change in the radius of curvature when connecting to the curve forming the back side at the point S 4 from the front edge of the arc shape. In order to prevent this sudden acceleration flow and deceleration flow on the blade back side, as shown in FIG. 7, on the downstream side of the straight line connecting the ventral end point P 3 and the leading edge arc center O,
It has been found from the inter-blade flow calculation that the end point S 3 on the back side may be located.
【００５４】尚以上の説明では翼前縁を正円の円弧状に
したものについて述べてきたが、この円弧は常に正円で
なければならないわけではなく、例えば楕円形であって
も同様な効果が得られるであろうし、またこの楕円形も
縦の楕円形でも横の楕円形でもほぼ同様な効果が得られ
るであろう。In the above description, the blade leading edge is made into a circular arc shape. However, this circular arc does not always have to be a perfect circle. The same effect can be obtained with this elliptical shape as well as a vertical elliptical shape or a horizontal elliptical shape.
【００５５】[0055]
【発明の効果】以上説明してきたように本発明は、翼前
縁側の横断面形状を円弧状にするとともに、翼の最大厚
み部分がこの円弧の中心部分に位置するようになしたか
ら、主流ガスが、翼の腹側と滑らかに接続した前縁円弧
上の一端点部分および翼の背側と滑らかに接続した前縁
円弧上の一端点部分に沿って流れるため、前縁部の急激
な高温ガスの加速が抑制され、翼表面の高温ガスの流速
を小さくすることができるので、ガス側の熱伝達率が下
がり、したがって翼内部を流通させる冷却空気の量を少
なくすることができる。As described above, according to the present invention, the cross-sectional shape of the leading edge side of the blade is arcuate, and the maximum thickness portion of the blade is located at the center of this arc. The gas flows along the one-point portion on the leading-edge arc that smoothly connects with the ventral side of the wing and the one-point portion on the leading-edge arc that smoothly connects with the back side of the wing. Since the acceleration of the hot gas is suppressed and the flow velocity of the hot gas on the blade surface can be reduced, the heat transfer coefficient on the gas side is reduced, and therefore the amount of cooling air flowing inside the blade can be reduced.
【図１】本発明の一実施例を示すものにしてその静翼を
示す横断面図。FIG. 1 is a cross-sectional view showing a vane of an embodiment of the present invention.
【図２】本発明静翼の線図。FIG. 2 is a diagram of a vane of the present invention.
【図３】本発明静翼を含むガスタービンの一部破断側面
図。FIG. 3 is a partially cutaway side view of a gas turbine including a vane of the present invention.
【図４】本発明静翼を含む段落装置の断面図。FIG. 4 is a sectional view of a paragraph device including a vane of the present invention.
【図５】静翼の線図。FIG. 5 is a diagram of a stationary blade.
【図６】静翼の線図。FIG. 6 is a diagram of a stationary blade.
【図７】静翼の線図。FIG. 7 is a diagram of a stationary blade.
【図８】本発明を適用した翼の翼面マッハ数分布を示す
図。FIG. 8 is a diagram showing a blade surface Mach number distribution of a blade to which the present invention is applied.
３…回転子、４…動翼、５…圧縮機、８…燃焼器、９…
静翼。3 ... Rotor, 4 ... Blade, 5 ... Compressor, 8 ... Combustor, 9 ...
Shizuka.
フロントページの続き (72)発明者 川池 和彦 茨城県土浦市神立町502番地 株式会社 日立製作所 機械研究所内 (72)発明者 竹原 勲 茨城県日立市幸町三丁目１番１号 株式 会社 日立製作所 日立工場内 (56)参考文献 特開 平３−138404（ＪＰ，Ａ) 特開 昭58−18565（ＪＰ，Ａ) 実開 昭63−191204（ＪＰ，Ｕ) 実開 昭49−98601（ＪＰ，Ｕ) 「タービン・発電機講座」、昭和53年 ４月発行、（社）火力発電技術協会Front Page Continuation (72) Inventor Kazuhiko Kawaike 502 Jinrachicho, Tsuchiura City, Ibaraki Prefecture Machinery Research Laboratory, Hitachi, Ltd. (72) Inventor Isao Takehara 3-1-1 Sachimachi Hitachi City, Ibaraki Hitachi Co., Ltd. Hitachi Ltd. In the factory (56) Reference JP-A-3-138404 (JP, A) JP-A-58-18565 (JP, A) Actually opened 63-191204 (JP, U) Actually opened 49-98601 (JP, U) ) "Turbine / Generator Course", published in April 1978, Thermal Power Generation Technology Association
Claims (9)
い徐々に大きくなり、かつその途中より後縁側に向かう
にしたがい徐々に小さくなるように形成され、かつ内部
に中空部を有し、該中空部に冷却媒体を流通せしめて翼
を内部から冷却するようになしたガスタービンの翼にお
いて、 前記翼前縁側の横断面形状を円弧状にするとともに、 該円弧の最大太り部分より後縁側に向かうにしたがいそ
の翼厚みを徐々に小さくするように該円弧の最大太り部
分から後縁までの領域を形成することを特徴とするガス
タービンの翼。1. A wing thickness is formed so that it gradually increases from the leading edge side toward the center side, and gradually decreases from that midpoint toward the trailing edge side, and has a hollow portion inside, In a blade of a gas turbine in which a cooling medium is circulated in the hollow portion to cool the blade from the inside, a cross-sectional shape of the leading edge side of the blade is an arc shape, and a trailing edge side is larger than a maximum thick portion of the arc. The thickest part of the arc so that the thickness of the blade gradually decreases toward
A blade of a gas turbine, which is characterized by forming a region from a minute to a trailing edge .
い徐々に大きくなり、かつその途中より後縁側に向かう
にしたがい徐々に小さくなるように形成され、かつ内部
に中空部を有し、該中空部に冷却媒体を流通せしめて翼
を内部から冷却するようになしたガスタービンの翼にお
いて、 前記翼前縁側の横断面形状を円弧状にするとともに、翼
の最大厚み部分がこの円弧の中心部分に位置するように
形成し、該円弧の最大厚み部分より後縁側に向かうにし
たがいその翼厚みを徐々に小さくするように該円弧の最
大厚み部分から後縁までの領域を形成することを特徴と
するガスタービンの翼。2. The blade is formed so that the blade thickness gradually increases from the leading edge side toward the center side and gradually decreases from the middle thereof toward the trailing edge side, and has a hollow portion inside, In a blade of a gas turbine in which a cooling medium is circulated in the hollow portion to cool the blade from the inside, the cross-sectional shape of the leading edge side of the blade is made into an arc shape, and the maximum thickness portion of the blade has this arc shape. It should be formed so that it is located in the central part, and it should be oriented toward the trailing edge from the maximum thickness part of the arc.
In order to gradually reduce the blade thickness,
A blade for a gas turbine, which is characterized by forming a region from a large thickness portion to a trailing edge .
い徐々に大きくなり、かつその途中より後縁側に向かう
にしたがい徐々に小さくなるように形成され、かつ内部
に中空部を有し、該中空部に冷却媒体を流通せしめて翼
を内部から冷却するようになしたガスタービンの翼にお
いて、 前記翼前縁側の横断面形状を略円弧状にするとともに、
該円弧の最大太り部分より後縁側に向かうにしたがい、
該円弧の最大太り部分から後縁に至る間の翼厚みを徐々
に小さくするようにしたことを特徴とするガスタービン
の翼。3. The blade is formed so that the blade thickness gradually increases from the leading edge side toward the center side and gradually decreases from the middle thereof toward the trailing edge side, and has a hollow portion inside, In a blade of a gas turbine, wherein a cooling medium is circulated in the hollow portion to cool the blade from the inside, the cross-sectional shape of the blade leading edge side is substantially arcuate,
As it goes toward the trailing edge side from the maximum fat part of the arc,
A blade of a gas turbine, characterized in that the blade thickness between the maximum thickened portion of the arc and the trailing edge is gradually reduced.
部より中央側に向かうにしたがい徐々に大きくなって最
大厚みを有し、かつこの最大厚みの部分より後縁側に向
かうにしたがい徐々に小さくなるように形成され、かつ
内部に中空部を有し、該中空部に冷却媒体を流通せしめ
て翼を内部から冷却するようになしたガスタービンの翼
において、 前記翼前縁側の円弧の直径を翼の最大厚みと等しく形成
し、該円弧の最大厚み部分より後縁側に向かうにしたが
い、該円弧の最大厚み部分から後縁に至る間の翼厚みを
徐々に小さくするようにしたことを特徴とするガスター
ビンの翼。4. A blade leading edge is formed in an arc shape, and the blade thickness gradually increases toward the center side from the arc portion so as to have a maximum thickness, and from the portion of this maximum thickness toward the trailing edge side. A blade of a gas turbine, which is formed so as to gradually decrease in size, has a hollow portion inside, and cools the blade from the inside by circulating a cooling medium in the hollow portion. equal form of arc diameter as the maximum thickness of the blade
However, the direction from the maximum thickness of the arc toward the trailing edge is
The blade thickness between the maximum thickness of the arc and the trailing edge
A blade of a gas turbine characterized by being gradually made smaller .
い徐々に大きくなり、かつその途中より後縁側に向かう
にしたがい徐々に小さくなるように形成され、かつ内部
に中空部を有し、該中空部に冷却媒体を流通せしめて翼
を内部から冷却するようになしたガスタービンの翼にお
いて、 前記翼前縁側の横断面形状を円弧状にするとともに、該
円弧の端部より後縁までの間の翼厚みを該円弧の端部よ
り後縁側に向かうにしたがい徐々に小さくするようにし
たことを特徴とするガスタービンの翼。5. The blade is formed so that the blade thickness gradually increases from the leading edge side toward the center side, and gradually decreases from the middle thereof toward the trailing edge side, and has a hollow portion inside, In a gas turbine blade in which a cooling medium is circulated in the hollow portion to cool the blade from the inside, a cross-sectional shape of the blade leading edge side is an arc shape, and
A blade of a gas turbine, wherein a blade thickness between the end of the arc and the trailing edge is gradually reduced toward the trailing edge side from the end of the arc.
い徐々に大きくなり、かつその途中より後縁側に向かう
にしたがい徐々に小さくなるように形成され、かつ内部
に中空部を有し、該中空部に冷却媒体を流通せしめて翼
を内部から冷却するようになしたガスタービンの翼にお
いて、 前記翼の横断面形状を、翼前縁側及び翼後縁側が夫々円
弧状に形成され、かつその両円弧端部同志が直線的に連
結された形に形成するようにしたことを特徴とするガス
タービンの翼。6. The blade is formed so that the blade thickness gradually increases from the leading edge side toward the center side, and gradually decreases from the middle thereof toward the trailing edge side, and has a hollow portion inside, In a blade of a gas turbine, wherein a cooling medium is circulated in the hollow portion to cool the blade from the inside, a cross-sectional shape of the blade is formed such that a blade leading edge side and a blade trailing edge side are each formed into an arc shape, and A blade of a gas turbine, characterized in that both arc ends thereof are linearly connected to each other.
い徐々に大きくなり、かつその途中より後縁側に向かう
にしたがい徐々に小さくなるように形成され、かつ内部
に中空部を有し、該中空部に冷却媒体を流通せしめて翼
を内部から冷却するようになしたガスタービンの翼にお
いて、 前記翼の横断面形状を、翼前縁側及び翼後縁側が夫々円
弧状に形成され、かつその両円弧端部同志が彎曲線にて
連結された形に形成するようにしたことを特徴とするガ
スタービンの翼。7. The blade thickness is gradually increased from the leading edge side toward the center side, and gradually reduced from the middle toward the trailing edge side, and has a hollow portion inside, In a blade of a gas turbine, wherein a cooling medium is circulated in the hollow portion to cool the blade from the inside, a cross-sectional shape of the blade is formed such that a blade leading edge side and a blade trailing edge side are each formed into an arc shape, and A blade of a gas turbine, characterized in that both arc end portions are formed so as to be connected by a curve.
い徐々に大きくなり、かつその途中より後縁側に向かう
にしたがい徐々に小さくなるように形成され、かつ内部
に中空部を有し、該中空部に冷却媒体を流通せしめて翼
を内部から冷却するようになしたガスタービンの翼にお
いて、前記翼前縁側の横断面形状を円弧状にするとともに、翼
の腹側部分と前記前縁側の円弧状の領域との端部から前
記円弧の中心を通り延長する直線は翼背側において前記
円弧状の領域に交わるよう構成されていることを特徴と
するガスタービンの翼。 8. A blade thickness is formed so as to gradually increase from the leading edge side toward the center side, and gradually decreases toward the trailing edge side from the middle thereof, and has a hollow portion inside, In a blade of a gas turbine in which a cooling medium is circulated in the hollow portion to cool the blade from the inside, the blade has a cross-sectional shape on the leading edge side that is an arc shape, and
From the end of the ventral part of the and the arcuate region on the front edge side
The straight line extending through the center of the circular arc is
Characterized by being configured to intersect with an arcuate region
Blades of a gas turbine to be used.
より冷却され、かつ内部から冷却される翼を有するガス
タービンにおいて、 前記翼を、その前縁側の横断面形状が円弧状に形成され
るとともに、該円弧の最大太り部分より後縁までの翼厚
みを該円弧の最大太り部分より後縁側に向かうにしたが
い徐々に小さくなるように形成したことを特徴とするガ
スタービン。 9. Compressed air from a compressor coupled to a turbine
Gas with blades that are more cooled and cooled from the inside
In the turbine, the blade is formed such that the cross-sectional shape on the leading edge side is arcuate.
And the blade thickness from the thickest part of the arc to the trailing edge
Although the direction toward the trailing edge from the maximum fat part of the arc is
The gas is characterized by being formed so that it gradually becomes smaller.
Sturbine.
Priority Applications (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP4249933A JP2684936B2 (en) | 1992-09-18 | 1992-09-18 | Gas turbine and gas turbine blade |
| US08/120,474 US5393198A (en) | 1992-09-18 | 1993-09-14 | Gas turbine and gas turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP4249933A JP2684936B2 (en) | 1992-09-18 | 1992-09-18 | Gas turbine and gas turbine blade |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH06101406A JPH06101406A (en) | 1994-04-12 |
| JP2684936B2 true JP2684936B2 (en) | 1997-12-03 |
Family
ID=17200336
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP4249933A Expired - Lifetime JP2684936B2 (en) | 1992-09-18 | 1992-09-18 | Gas turbine and gas turbine blade |
Country Status (2)
| Country | Link |
|---|---|
| US (1) | US5393198A (en) |
| JP (1) | JP2684936B2 (en) |
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| US5716192A (en) * | 1996-09-13 | 1998-02-10 | United Technologies Corporation | Cooling duct turn geometry for bowed airfoil |
| US6257830B1 (en) * | 1997-06-06 | 2001-07-10 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade |
| US6430931B1 (en) * | 1997-10-22 | 2002-08-13 | General Electric Company | Gas turbine in-line intercooler |
| JP2001355404A (en) * | 2000-06-15 | 2001-12-26 | Ishikawajima Harima Heavy Ind Co Ltd | Method of forming shape of turbine blade surface |
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| US6884036B2 (en) | 2003-04-15 | 2005-04-26 | General Electric Company | Complementary cooled turbine nozzle |
| ITMI20041804A1 (en) * | 2004-09-21 | 2004-12-21 | Nuovo Pignone Spa | SHOVEL OF A RUTOR OF A FIRST STAGE OF A GAS TURBINE |
| US7956102B2 (en) * | 2007-04-09 | 2011-06-07 | The Board Of Trustees Of The University Of Illinois | Sol-gel inks |
| CN102016234B (en) * | 2008-03-28 | 2015-05-20 | 阿尔斯通技术有限公司 | Guide vane for a gas turbine |
| ES2400197T3 (en) * | 2009-09-04 | 2013-04-08 | Alstom Technology Ltd | Gas turbine |
| EP2397653A1 (en) * | 2010-06-17 | 2011-12-21 | Siemens Aktiengesellschaft | Platform segment for supporting a nozzle guide vane for a gas turbine and method of cooling thereof |
| US10519781B2 (en) | 2017-01-12 | 2019-12-31 | United Technologies Corporation | Airfoil turn caps in gas turbine engines |
| US10465528B2 (en) | 2017-02-07 | 2019-11-05 | United Technologies Corporation | Airfoil turn caps in gas turbine engines |
| US10480329B2 (en) | 2017-04-25 | 2019-11-19 | United Technologies Corporation | Airfoil turn caps in gas turbine engines |
| US10267163B2 (en) * | 2017-05-02 | 2019-04-23 | United Technologies Corporation | Airfoil turn caps in gas turbine engines |
| DE102017212310A1 (en) * | 2017-07-19 | 2019-01-24 | MTU Aero Engines AG | Shovel, blade ring, blade ring segment and turbomachine |
| KR102048863B1 (en) | 2018-04-17 | 2019-11-26 | 두산중공업 주식회사 | Turbine vane having insert supports |
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| US1777098A (en) * | 1927-06-30 | 1930-09-30 | Ljungstroms Angturbin Ab | Blade system of gas or steam turbines |
| US2830753A (en) * | 1951-11-10 | 1958-04-15 | Edward A Stalker | Axial flow compressors with circular arc blades |
| US2866618A (en) * | 1953-02-13 | 1958-12-30 | Thomas W Jackson | Reverse flow air cooled turbine blade |
| US3623318A (en) * | 1970-06-29 | 1971-11-30 | Avco Corp | Turbine nozzle cooling |
| GB1560683A (en) * | 1972-11-28 | 1980-02-06 | Rolls Royce | Turbine blade |
| JPS4998601U (en) * | 1972-12-15 | 1974-08-26 | ||
| FR2474095B1 (en) * | 1980-01-17 | 1986-02-28 | Rolls Royce | VIBRATION DAMPING DEVICE FOR MOBILE BLADES OF A GAS TURBINE ENGINE |
| DE3110098C2 (en) * | 1981-03-16 | 1983-03-17 | MTU Motoren- und Turbinen-Union München GmbH, 8000 München | Turbine guide vane for gas turbine engines |
| JPS5818565A (en) * | 1981-07-27 | 1983-02-03 | Tsuchiya Riyuuko | Universal turbine |
| JPS63191204U (en) * | 1987-05-29 | 1988-12-09 | ||
| JP2716480B2 (en) * | 1988-10-18 | 1998-02-18 | 西脇 邦彦 | Gas turbine blade cooling method |
| US4962640A (en) * | 1989-02-06 | 1990-10-16 | Westinghouse Electric Corp. | Apparatus and method for cooling a gas turbine vane |
| JP3142850B2 (en) * | 1989-03-13 | 2001-03-07 | 株式会社東芝 | Turbine cooling blades and combined power plants |
| US5035578A (en) * | 1989-10-16 | 1991-07-30 | Westinghouse Electric Corp. | Blading for reaction turbine blade row |
| US5281084A (en) * | 1990-07-13 | 1994-01-25 | General Electric Company | Curved film cooling holes for gas turbine engine vanes |
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1992
- 1992-09-18 JP JP4249933A patent/JP2684936B2/en not_active Expired - Lifetime
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1993
- 1993-09-14 US US08/120,474 patent/US5393198A/en not_active Expired - Lifetime
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| Title |
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| 「タービン・発電機講座」、昭和５３年４月発行、（社）火力発電技術協会 |
Also Published As
| Publication number | Publication date |
|---|---|
| JPH06101406A (en) | 1994-04-12 |
| US5393198A (en) | 1995-02-28 |
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